(1) Field of the Invention
The present invention relates to a method of controlling movable pitching stabilizer means, and to a stabilizer system and to an aircraft applying the method. More particularly, the invention lies in the narrow technical field of means for stabilizing rotorcraft in pitching.
(2) Description of Related Art
Conventionally, a rotorcraft comprises, by way of example, a fuselage extending longitudinally from a front end to a rear end on either side of an anteroposterior plane of symmetry, and in a vertical direction from a bottom portion having landing gear to a top portion having a rotary wing. The rotary wing may include at least one main rotor for providing lift and possibly also propulsion.
Furthermore, the rotorcraft may have a tail rotor at its rear end. The tail rotor serves in particular to control yaw movements of the rotorcraft.
Furthermore, a rotorcraft sometimes has additional stabilizer surfaces. For example, a rotorcraft is usually fitted with a surface for stabilizing yaw movements.
Such a surface for stabilizing yaw movements is generally referred to as a “fin”.
Likewise, a rotorcraft sometimes has means for balancing and stabilizing pitching movements, referred to more simply as “stabilizer” means. Stabilizer means present an angle of absolute value lying in the range 0° to substantially 90° relative to said anteroposterior plane. The pitching stabilizer means may possibly comprise two pitching stabilizer surfaces extending symmetrically on either side of an anteroposterior plane of symmetry of the rotorcraft, being orthogonal to said anteroposterior plane, or indeed presenting a V-shape, for example.
Such means for stabilizing pitching movements are sometimes referred to as a “horizontal stabilizer” or more simply below as a “stabilizer”. The term “stabilizer” is used more particularly since the stabilizer means are not necessarily horizontal. The term “pitching stabilizer means” is also used.
A stabilizer may comprise at least one airfoil surface passing through the rear end of the aircraft in a transverse direction, or at least one non-through airfoil surface extending transversely from said rear end on one side only of said anteroposterior plane.
In hovering flight, balancing the pitching moment of the rotorcraft about a fixed point involves two main components: a first component due to the weight and the position of the center of gravity of the rotorcraft, and a second component due to the aerodynamic forces resulting from the main lift rotor. For a given weight of the rotorcraft, the second component is proportional to the angle of tilt of the main rotor relative to a vertical direction. Furthermore, variations in the position of the center of gravity of the rotorcraft give rise to variations in the attitude of the rotorcraft.
While the rotorcraft is in cruising flight, a third component of the pitching moment becomes involved: the aerodynamic moment that is exerted on the fuselage of the rotorcraft as a result of variations in the angle of incidence of the fuselage relative to the upstream air flow. The aerodynamic pitching moment tends to move the aircraft away from its equilibrium position. This third component, which is unstable, has the consequence of increasing variations in the longitudinal attitude that are associated with the position of the center of gravity relative to the variations that are observed during hovering flight.
These variations in attitude have negative consequences. Considerable nose-down attitudes increase the aerodynamic drag of the rotorcraft and consequently reduce its maximum speed. Such nose-down attitudes also lead to a feeling of discomfort for the crew and passengers. Considerable nose-up attitudes give rise to large moments on the drive mast of the main rotor and on the hub of the main rotor, with consequences that are unfavorable for the lifetimes of those elements. At lower speeds of advance and during stages of landing, large nose-up attitudes also give rise to a reduction in the pilot's visibility and thus to an increase in the pilot's workload.
The pitching stabilizer means placed towards the rear of a rotorcraft seek in particular to compensate for the instability of the pitching moment of the fuselage and to balance the attitude of the rotorcraft. It is difficult to dimension pitching stabilizer means.
In order to optimize the performance of an aircraft at high speed, the pitching stabilizer means are dimensioned so as to obtain a longitudinal attitude that is close to a zero attitude. Such a dimensioning is nevertheless penalizing on the operation of the main rotor, given that a considerable nose-down attitude is desirable for the operation of the main rotor when the forward speed of the rotorcraft is high.
Furthermore, the dimensioning needs to be satisfactory for the various potential configurations of the weight, the attitude, the position of the center of gravity of the aircraft, and possibly also various possible external aerodynamic configurations of the aircraft.
Optionally, the effectiveness of the pitching stabilizer means can be maximized by increasing the wing area so as to reduce the effects of disturbances associated with variations in the weight and the positioning of the center of gravity of the rotorcraft.
Nevertheless, such a solution is limited, e.g. because of the “attitude hump” phenomenon that is known to the person skilled in the art and that results from interactions between the main rotor and the pitching stabilizer means. Furthermore, a large wing area tends in particular to lead to large variations in the attitude of the aircraft during stages of flight in which the stream of air coming from the main rotor impacts against the pitching stabilizer means, e.g. while the aircraft is climbing or descending.
Manufacturers have sought to remedy those drawbacks by creating a device referred to for convenience as a “positioner” device. The function of such a positioner device is to adjust the angular position of pitching stabilizer means so as to balance the rotorcraft in pitching, while simultaneously controlling its performance and the loads applied to the hub of a main rotor, while being unaffected by variations in the position of the center of gravity and while satisfying the above-mentioned constraints that require the size of said aerodynamic surface to be limited.
Thus, the pitching stabilizer means are movable in pivoting about a pivot axis. As a result, the positioner device enables a deflection angle of the stabilizer means to be adjusted. Such a deflection angle represents an angle between a reference chord of the pitching stabilizer means and a reference plane of the aircraft, e.g. a horizontal reference plane.
Thus, Documents U.S. Pat. No. 2,424,882 and GB 657 796 provide a positioner device comprising a lever that is mechanically connected to pitching stabilizer means in order to control the position of the stabilizer means.
Those documents suggest manual piloting of the pitching stabilizer means. The way the pitching stabilizer means are controlled then represents a compromise that is implemented directly by the pilot of the aircraft.
Likewise, Document WO 2004/007282 A2 describes means for controlling a deflectable stabilizer independently of a main cyclic control of a helicopter. Those control means comprise a control that can be moved by a pilot.
Given the increasing technical complexity of rotorcraft and of piloting them, pitching stabilizer means are controlled while taking account of various factors, rather than merely on the basis of an action by the pilot.
Document FR 2 456 663 provides for using a positioner device to servo-control the position of pitching stabilizer means as a function of flight conditions.
That positioner device comprises an electric motor for servo-controlling the control of the pitching stabilizer means. Furthermore, a sensor sensing the longitudinal cyclic control, a sensor sensing the collective pitch position of the blades of the main rotor, and a sensor for sensing speed relative to air issue signals that are combined in order to provide a sum signal. The sum signal is modified by a factor that varies inversely with the speed of the aircraft relative to air, and it is then transmitted to a servo-control device for controlling the pitching stabilizer means.
In order to operate pitching stabilizer means, certain positioner devices make use not of one electric actuator, but rather of two that are mounted back to back. One actuator is fastened to the tail boom of said aircraft, and another actuator is fastened to pitching stabilizer means.
Document EP 1 547 920 A1 describes a tiltable stabilizer seeking to reduce the vibration that is generated on the structure of a helicopter by the aerodynamic flow of air coming from the main rotor. That document makes use of a positioner device having at least one vibration sensor.
Document WO 2008/142256 suggests adjusting the pitching moment of a fuselage from at least one movable pitching control surface forming part of the horizontal stabilizer of an aircraft. A one-to-one relationship may be established between the position of the stabilizer and the value of the moment exerted on a rotor mast. The value of this moment can thus be controlled at all times firstly with an objective of reducing fatigue on the mechanical part and secondly with an objective of sharing power between the propellers and a rotor of the aircraft.
Document FR 2 383 475 describes a helicopter having pitching stabilizer means and a positioner device for modifying the position of the pitching stabilizer means.
Under such circumstances, the aircraft has generator means for transmitting a control signal to the positioner device. The control signal is generated as a function of an aerodynamic speed, of a collective pitch percentage of the blades of a rotor, of lateral accelerations of the helicopter, of a pitching rate of the about a pitching axis, and of a constant produced by a bias circuit.
Document FR 2 067 224 describes a mechanical positioner device actuated by inertia and suitable for causing a stabilizer to pivot.
Document FR 2 962 972 is remote from the problem of the invention since it presents stationary biplane stabilizer means.
Documents FR 2 771 706, WO 2015/152910, EP 0 183 282, and EP 0 743 582 are also known.